BACKGROUND
[0001] This disclosure relates to gas turbine engines and, more particularly, to an engine
having a geared turbo fan architecture that is designed to efficiently operate with
a high bypass ratio and a low pressure ratio.
[0002] The propulsive efficiency of a gas turbine engine depends on many different factors,
such as the design of the engine and the resulting performance debits on the fan that
propels the engine. As an example, the fan rotates at a high rate of speed such that
air passes over the blades at transonic or supersonic speeds. The fast-moving air
creates flow discontinuities or shocks that result in irreversible propulsive losses.
Additionally, physical interaction between the fan and the air causes downstream turbulence
and further losses. Although some basic principles behind such losses are understood,
identifying and changing appropriate design factors to reduce such losses for a given
engine architecture has proven to be a complex and elusive task.
SUMMARY
[0003] An exemplary gas turbine engine includes a spool, a turbine coupled to drive the
spool, and a propulsor that is coupled to be driven by the turbine through the spool.
A gear assembly is coupled between the propulsor and the spool such that rotation
of the turbine drives the propulsor at a different speed than the spool. The propulsor
includes a hub and a row of propulsor blades that extend from the hub. The row includes
no more than 20 of the propulsor blades.
[0004] In another aspect, a gas turbine engine includes a core flow passage and a bypass
flow passage. A propulsor is arranged at an inlet of the bypass flow passage and core
flow passage. The propulsor includes a hub and a row of propulsor blades that extend
from the hub. The row includes no more than 20 of the propulsor blades and the bypass
flow passage has a design pressure ratio of approximately 1.3-1.55 with regard to
an inlet pressure and an outlet pressure of the bypass flow passage.
[0005] An exemplary propulsor for use in a gas turbine engine includes a rotor having a
row of propulsor blades that extends radially outwardly from a hub. Each of the propulsor
blades extends radially between a root and a tip and in a chord direction between
a leading edge and a trailing edge to define a chord dimension at the tip of each
propulsor blade. The row of propulsor blades defines a circumferential pitch with
regard to the tips. The row of propulsor blades has a solidity value defined as the
chord dimension divided by the circumferential pitch. The row also includes a number
of the propulsor blades that is no greater than 20 such that a ratio of the number
of propulsor blades to the solidity value is from 9 to 20.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] The various features and advantages of the disclosed examples will become apparent
to those skilled in the art from the following detailed description. The drawings
that accompany the detailed description can be briefly described as follows.
Figure 1 is a schematic cross-section of a gas turbine engine.
Figure 2 is a perspective view of a fan section of the engine of Figure 1.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
[0007] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine
20 may be a two-spool turbofan that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28. Alternative engine architectures
may include a single-spool design, a three-spool design, or an open rotor design,
among other systems or features.
[0008] The fan section 22 drives air along a bypass flow passage B while the compressor
section 24 drives air along a core flow passage C for compression and communication
into the combustor section 26. Although depicted as a turbofan gas turbine engine,
it is to be understood that the concepts described herein are not limited to use with
turbofans and the teachings may be applied to other types of gas turbine engines.
[0009] The engine 20 includes a low speed spool 30 and high speed spool 32 mounted for rotation
about an engine central longitudinal axis A relative to an engine static structure
36 via several bearing systems 38. The low speed spool 30 generally includes an inner
shaft 40 that is coupled with a propulsor 42, a low pressure compressor 44 and a low
pressure turbine 46. The low pressure turbine 46 drives the propulsor 42 through the
inner shaft 40 and a gear assembly 48, which allows the low speed spool 30 to drive
the propulsor 42 at a different (e.g. lower) angular speed.
[0010] The high speed spool 32 includes an outer shaft 50 that is coupled with a high pressure
compressor 52 and a high pressure turbine 54. A combustor 56 is arranged between the
high pressure compressor 52 and the high pressure turbine 54. The inner shaft 40 and
the outer shaft 50 are concentric and rotate about the engine central longitudinal
axis A, which is collinear with their longitudinal axes.
[0011] A core airflow in core flow passage C is compressed by the low pressure compressor
44 then the high pressure compressor 52, mixed with the fuel in the combustor 56,
and then expanded over the high pressure turbine 54 and low pressure turbine 46. The
turbines 54, 46 rotationally drive the respective low speed spool 30 and high speed
spool 32 in response to the expansion.
[0012] As shown, the propulsor 42 is arranged at an inlet 60 of the bypass flow passage
B and core flow passage C. Air flow through the bypass flow passage B exits the engine
20 through an outlet 62 or nozzle. For a given design of the propulsor 42, the inlet
60 and the outlet 62 of the engine 20 define a design pressure ratio with regard to
an inlet pressure at the inlet 60 and an outlet pressure at the outlet 62 of the bypass
flow passage B. As an example, the design pressure ratio may be determined based upon
the stagnation inlet pressure and the stagnation outlet pressure at a design rotational
speed of the engine 20. In that regard, the engine 20 may optionally include a variable
area nozzle 64 within the bypass flow passage B. The variable area nozzle 64 is operative
to change a cross-sectional area 66 of the outlet 62 to thereby control the pressure
ratio via changing pressure within the bypass flow passage B. The design pressure
ratio may be defined with the variable area nozzle 64 fully open or fully closed.
[0013] Referring to Figure 2, the propulsor 42, which in this example is a fan, includes
a rotor 70 having a row 72 of propulsor blades 74 that extend a circumferentially
around a hub 76. Each of the propulsor blades 74 extends radially outwardly from the
hub 76 between a root 78 and a tip 80 and in a chord direction (axially and circumferentially)
between a leading edge 82 and a trailing edge 84. A chord dimension (CD) is a length
between the leading edge 82 and the trailing edge 84 at the tip of each propulsor
blade 74. The row 72 of propulsor blades 74 also defines a circumferential pitch (CP)
that is equivalent to the arc distance between the tips 80 of neighboring propulsor
blades 74.
[0014] As will be described, the example propulsor 42 includes a number (N) of the propulsor
blades 74 and a geometry that, in combination with the architecture of the engine
20, provides enhanced propulsive efficiency by reducing performance debits of the
propulsor 42.
[0015] In the illustrated example, the number N of propulsor blades in the row 72 is no
more than 20. In one example, the propulsor 42 includes 18 of the propulsor blades
74 uniformly circumferentially arranged about the hub 76. In other embodiments, the
number N may be any number of blades from 12-20.
[0016] The propulsor blades 74 define a solidity value with regard to the chord dimension
CD and the circumferential pitch CP. The solidity value is defined as a ratio (R)
of CD/CP (i.e., CD divided by CP). In embodiments, the solidity value of the propulsor
42 is between 1.0 and 1.3. In further embodiments, the solidity value is from 1.1
to 1.2.
[0017] Additionally, in combination with the given example solidity values, the engine 20
may be designed with a particular design pressure ratio. In embodiments, the design
pressure ratio may be between 1.3 and 1.55. In a further embodiment, the design pressure
ratio may be between 1.3 and 1.4.
[0018] The engine 20 may also be designed with a particular bypass ratio with regard to
the amount of air that passes through the bypass flow passage B and the amount of
air that passes through the core flow passage C. As an example, the design bypass
ratio of the engine 20 may nominally be 12, or alternatively in a range of approximately
8.5 to 13.5.
[0019] The propulsor 42 also defines a ratio of N/R. In embodiments, the ratio N/R is from
9 to 20. In further embodiments, the ratio N/R is from 14 to 16. The table below shows
additional examples of solidity and the ratio N/R for different numbers of propulsor
blades 74.
TABLE: Number of Blades, Solidity and Ratio N/R
Number of Blades (N) |
Solidity |
Ratio N/R |
20 |
1.3 |
15.4 |
18 |
1.3 |
13.8 |
16 |
1.3 |
12.3 |
14 |
1.3 |
10.8 |
12 |
1.3 |
9.2 |
20 |
1.2 |
16.7 |
18 |
1.2 |
15.0 |
16 |
1.2 |
13.3 |
14 |
1.2 |
11.7 |
12 |
1.2 |
10.0 |
20 |
1.1 |
18.2 |
18 |
1.1 |
16.4 |
16 |
1.1 |
14.5 |
14 |
1.1 |
12.7 |
12 |
1.1 |
10.9 |
20 |
1.0 |
20.0 |
18 |
1.0 |
18.0 |
16 |
1.0 |
16.0 |
14 |
1.0 |
14.0 |
12 |
1.0 |
12.0 |
[0020] The disclosed ratios of N/R enhance the propulsive efficiency of the disclosed engine
20. For instance, the disclosed ratios of N/R are designed for the geared turbo fan
architecture of the engine 20 that utilizes the gear assembly 48. That is, the gear
assembly 48 allows the propulsor 42 to rotate at a different, lower speed than the
low speed spool 30. In combination with the variable area nozzle 64, the propulsor
42 can be designed with a large diameter and rotate at a relatively slow speed with
regard to the low speed spool 30. A relatively low speed, relatively large diameter,
and the geometry that permits the disclosed ratios of N/R contribute to the reduction
of performance debits, such as by lowering the speed of the air or fluid that passes
over the propulsor blades 74.
[0021] Although a combination of features is shown in the illustrated examples, not all
of them need to be combined to realize the benefits of various embodiments of this
disclosure. In other words, a system designed according to an embodiment of this disclosure
will not necessarily include all of the features shown in any one of the Figures or
all of the portions schematically shown in the Figures. Moreover, selected features
of one example embodiment may be combined with selected features of other example
embodiments.
[0022] The preceding description is exemplary rather than limiting in nature. Variations
and modifications to the disclosed examples may become apparent to those skilled in
the art that do not necessarily depart from the essence of this disclosure. The scope
of legal protection given to this disclosure can only be determined by studying the
following claims.
1. A gas turbine engine (20) comprising:
a spool (30);
a turbine (46) coupled to drive the spool (30);
a propulsor (42) coupled to be driven by said turbine (46) through said spool (30);
and
a gear assembly (48) coupled between said propulsor (42) and said spool (30) such
that rotation of said spool (30) drives said propulsor (42) at a different speed than
said spool (30),
wherein said propulsor (42) includes a hub (76) and a row (72) of propulsor blades
(74) that extend from said hub (76), and said row (72) includes a number (N) of said
propulsor blades (74) that is no more than 20.
2. The gas turbine engine as recited in claim 1, wherein said propulsor (42) is located
at an inlet of a bypass flow passage (B) having a design pressure ratio that is between
1.3 and 1.55 with regard to an inlet pressure and an outlet pressure of said bypass
flow passage.
3. A gas turbine engine (20) comprising:
a core flow passage (C);
a bypass flow passage (B); and
a propulsor (42) arranged at an inlet of said bypass flow passage (B) and said core
flow passage (C), said propulsor (42) including a hub (76) and a row (72) of propulsor
blades that extend from said hub (76), and said row (72) includes no more than 20
of said propulsor blades (74), and
wherein said bypass flow passage (B) has a design pressure ratio of 1.3 - 1.55 with
regard to an inlet pressure and an outlet pressure of said bypass flow passage (B).
4. The gas turbine engine as recited in claim 3, wherein said propulsor (42) is coupled
to be driven by a turbine (46) through a spool (30), and a gear assembly (48) is coupled
between said propulsor (42) and said spool (30) such that rotation of said turbine
(46) drives said propulsor (42) at a different speed than said spool (30).
5. The gas turbine engine as recited in claim 2, 3 or 4 wherein said design pressure
ratio is between 1.3 and 1.4.
6. The gas turbine engine as recited in any preceding claim, wherein each of said propulsor
blades (74) extends radially between a root (78) and a tip (80) and in a chord direction
between a leading edge (82) and a trailing edge (84) at the tip (80) to define a chord
dimension (CD), said row (72) of propulsor blades (74) defining a circumferential
pitch (CP) with regard to said tips (80), wherein said row (72) of propulsor blades
(74) has a solidity value (R) of CD/CP that is between 1.0 and 1.3.
7. The gas turbine engine as recited in claim 6, wherein R is from 1.1 to 1.2.
8. The gas turbine as recited in claim 6 or 7, wherein a ratio of N/R is from 9 to 20.
9. The gas turbine as recited in claim 8, wherein said ratio of N/R is 14 to 16.
10. The gas turbine engine as recited in any preceding claim, wherein said propulsor (42)
is located at an inlet of a core flow passage (C) and a bypass flow passage (B) that
define a design bypass ratio of approximately 12 with regard to flow through said
core flow passage (C) and said bypass flow passage (B).
11. The gas turbine engine as recited in claim 10, wherein said propulsor (42) and said
bypass flow passage (B) define a design pressure ratio that is between 1.3 and 1.55.
12. The gas turbine engine as recited in claim 11, wherein said design pressure ratio
is between 1.3 and 1.4.
13. The gas turbine engine as recited in any preceding claim, further comprising a low
pressure compressor section (44) and a low pressure turbine section (46) that are
each coupled to be driven through said spool (30), and a high pressure compressor
section (52) and a high pressure turbine section (54) that are coupled to be driven
through another spool (32).
14. A propulsor (42) for use in a gas turbine engine, the propulsor (42) comprising:
a rotor including a row (72) of propulsor blades (74) extending radially outwardly
from a hub (76), each of said propulsor blades (74) extending radially between a root
(78) and a tip (80) and in a chord direction between a leading edge (82) and a trailing
edge (84) to define a chord dimension (CD) at the tip (80) of each propulsor blade
(74), said row (72) of propulsor blades (74) defining a circumferential pitch (CP)
with regard to the tips (80),
wherein said row (72) of propulsor blades has a solidity value (R) defined as CD/CP,
and said row (72) includes a number (N) of said propulsor blades (74) that is no greater
than 20 such that a ratio ofN/R is from 9 to 20.
15. The propulsor as recited in claim 14, wherein said ratio of N/R is 14 to 16.